Turbocharger with mixed flow turbine stage

ABSTRACT

A turbocharger is disclosed for use with an engine. The turbocharger may include a housing at least partially defining a compressor shroud and a turbine shroud. The turbocharger may also include a compressor wheel disposed within the compressor shroud, a shaft connected to the compressor wheel, and a turbine wheel disposed within the turbine shroud and connected to an end of the shaft opposite the compressor wheel. The turbine wheel may have a generally annular hub, and a plurality of blades disposed radially around the hub. Each of the plurality of blades may include an airfoil having a hub face connected to the hub, a shroud face opposite the hub face and oriented towards the turbine shroud, a trailing edge, and a leading edge opposite the trailing edge. The leading edge may be substantially straight or substantially concave in a meridional plane. An angle between a base of the hub and the leading edge may be about 25-55 degrees. The turbocharger may also include a nozzle ring having a ring-shaped generally flat plate located at a periphery of the turbine wheel, and a plurality of vanes disposed radially around an upper surface of the plate. A camber of each of the plurality of vanes may be generally S-shaped along its meridional length from a leading edge to a trailing edge of each of the plurality of vanes.

TECHNICAL FIELD

The present disclosure is directed to a turbocharger and, more particularly, to a turbocharger with a mixed flow turbine stage.

BACKGROUND

Internal combustion engines such as, for example, diesel engines, gasoline engines, and gaseous fuel powered engines are supplied with a mixture of air and fuel for subsequent combustion within the engines that generates a mechanical power output. In order to increase the power output generated by this combustion process, an engine can be equipped with a turbocharged air induction system.

A turbocharged air induction system includes a turbocharger that uses exhaust from the engine to compress air flowing into the engine, thereby forcing more air into a combustion chamber of the engine than the engine could otherwise draw into the combustion chamber. This increased supply of air allows for increased fueling, resulting in an increased power output. A turbocharged engine typically produces more power than the same engine without turbocharging.

A conventional turbocharger includes a turbine housing, and a turbine wheel centrally disposed within the housing and driven by exhaust to rotate a connected compressor wheel. The exhaust is pushed against blades connected to the turbine wheel to cause rotation of the turbine wheel. In some applications, vanes disposed on a nozzle ring connected to the turbine wheel accelerate the exhaust through the blades. The vanes and/or the blades of the turbine can direct the exhaust in axial, radial, and tangential directions.

A mixed flow turbine is generally viewed as across design between a radial and an axial turbine. An exemplary mixed flow turbine is disclosed in U.S. Pat. No. 8,128,356 to Higashimori that issued on Mar. 6, 2012 (the '356 patent). Specifically, the '356 patent describes a mixed flow turbine having blades whose outline of leading edges located at an upstream side is formed in a convex shape toward the upstream side, and a scroll that is a space formed upstream of the blades by a casing having a shroud that covers the radially external edges of the blades. Working fluid is supplied at a hub and the shroud and flows substantially in axial, radial, and tangential directions at a shroud-side inlet channel and at a hub-side inlet channel. A shape of the leading edges of the blades is designed to reduce incidence loss.

Although the mixed flow turbine of the '356 patent may be adequate for some applications, it may still be less than optimal at wide operating conditions. In particular, the mixed flow turbine of the '356 patent directs a non-uniform and poorly guided mixed flow through the turbine stage at wide operating conditions, which can result in high energy losses, reduced aerodynamic efficiencies, and increased mechanical or vibrational stresses (or strains) on the turbine during operation due to flow misalignment (high incidence) with the blades of the turbine. Also, the blade angle and thickness distributions of the mixed flow turbine shown in '356 patent are generally not smooth like a Bezier curve, which can lead to problems manufacturing the blades.

The turbocharger of the present disclosure solves one or more of the problems set forth above and/or other problems of the prior art.

SUMMARY

In one aspect, the present disclosure is directed to a turbocharger. The turbocharger may include a housing at least partially defining a compressor shroud and a turbine shroud. The turbocharger may also include a compressor wheel disposed within the compressor shroud, a shaft connected to the compressor wheel, and a turbine wheel disposed within the turbine shroud and connected to an end of the shaft opposite the compressor wheel. The turbine wheel may have a generally annular hub, and a plurality of blades disposed radially around the hub. Each of the plurality of blades may include an airfoil having a hub face connected to the hub, a shroud face opposite the hub face and oriented towards the turbine shroud, a trailing edge, and a leading edge opposite the trailing edge. An angle between a base of the hub and the leading edge may be about 25-55 degrees. The leading edge may be substantially straight or substantially concave in a meridional plane. The turbocharger may also include a nozzle ring having a ring-shaped generally flat plate located at a periphery of the turbine wheel, and a plurality of vanes disposed radially around an upper surface of the plate. A camber of each of the plurality of vanes may be generally S-shaped along its meridional length from a leading edge to a trailing edge of each of the plurality of vanes.

In a second aspect, the present disclosure is directed to a turbine blade for a turbocharger. The turbine blade may include an airfoil having a hub face connected to a turbine wheel hub of the turbocharger, a shroud face located opposite the hub face and oriented towards a turbine shroud of the turbocharger, a trailing edge, and a leading edge opposite the trailing edge. An angle between a base of the turbine wheel and the leading edge may be about 25-55 degrees. The leading edge may be substantially straight or substantially concave in a meridional plane.

In a third aspect, the present disclosure is directed to nozzle ring for a turbocharger. The nozzle ring may include a ring-shaped generally flat plate having an inner annular hub, and an outer annular flange radially spaced apart from the inner annular hub. The nozzle ring may also include a plurality of vanes disposed between the inner annular hub and the outer annular flange. A camber of each of the plurality of vanes may be generally S-shaped along its meridional length from a leading edge to a trailing edge of each of the plurality of vanes.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic illustration of an exemplary disclosed power system;

FIG. 2 is a cross-sectional illustration of an exemplary disclosed turbocharger that may be used in conjunction with the power system of Ea. 1;

FIG. 3 is a pictorial illustration of an exemplary disclosed turbine wheel and nozzle ring that may be used in conjunction with the turbocharger of FIG. 2;

FIG. 4 is a side-view illustration of the turbine wheel of FIG. 3;

FIG. 5 is a meridional-view illustration of an exemplary disclosed turbine blade that may be used in conjunction with the turbine wheel of FIG. 4;

FIG. 6 is a meridional-view illustration of an alternative embodiment of the turbine blade of FIG. 5;

FIGS. 7 and 8 are charts associated with exemplary disclosed geometry of the turbine blade of FIGS. 5; and

FIGS. 9, 10, and 11 are charts associated with exemplary disclosed geometry of the nozzle ring of FIG. 3.

DETAILED DESCRIPTION

FIG. 1 illustrates a power system 10 having an engine 12, an air induction system 14, and an exhaust system 16. For the purposes of this disclosure, engine 12 is depicted and described as a four-stroke diesel engine. One skilled in the art will recognize, however, that engine 12 may be any other type of combustion engine such as, for example, a two- or four-stroke gasoline or gaseous fuel-powered engine. Air induction system 14 may be configured to direct air or a mixture of air and fuel into engine 12 for combustion. Exhaust system 16 may be configured to direct combustion exhaust from engine 12 to the atmosphere.

Engine 12 may include an engine block 18 that at least partially defines a plurality of cylinders 20. A piston (not shown) may be slidably disposed within each cylinder 20 to reciprocate between a top-dead-center position and a bottom-dead-center position, and a cylinder head (not shown) may be associated with each cylinder 20. Each cylinder 20, piston, and cylinder head may together at least partially define a combustion chamber. In the illustrated embodiment, engine 12 includes twelve cylinders 20 arranged in a V-configuration (i.e., a configuration having first and second banks 22, 24 or rows of cylinders 20). However, it is contemplated that engine 12 may include a greater or lesser number of cylinders 20 and that cylinders 20 may be arranged in an inline configuration, in an opposing-piston configuration, or in another configuration, as desired.

Air induction system 14 may include, among other things, at least one compressor 28 that may embody a fixed geometry compressor, a variable geometry compressor, or any other type of compressor configured to receive air and compress the air to a desired pressure level. Compressor 28 may direct air to one or more intake manifolds 30 associated with engine 12. It should be noted that air induction system 14 may include multiple compressors 28 arranged in a serial configuration, a parallel configuration, or a combination serial/parallel configuration.

Exhaust system 16 may include, among other things, an exhaust manifold 34 connected to one or both of banks 22, 24 of cylinders 20. Exhaust system 16 may also include at least one turbine 32 driven by the exhaust from exhaust manifold 34 to rotate compressor 28 of air induction system 14. Compressor 28 and turbine 32 may together form a turbocharger 36. Turbine 32 may be configured to receive exhaust and convert potential energy in the exhaust to a mechanical rotation. After exiting turbine 32, the exhaust may be discharged to the atmosphere through an aftertreatment system 38 that may include, for example, a hydrocarbon closer, a diesel oxidation catalyst (DOC), a diesel particulate filter (DPF), and/or any other treatment device known in the art, if desired. It should be noted that exhaust system 16 may include multiple turbines 32 arranged in a serial configuration, a parallel configuration, or a combination serial/parallel configuration, as desired.

As illustrated in FIG. 2, compressor 28 and turbine 32 of turbocharger 36 may be connected to each other via a common shaft 50. Turbocharger 36 may include a housing 40 at least partially defining compressor and turbine shrouds 42, 44 that are configured to house corresponding compressor and turbine wheels 46, 48. Compressor shroud 42 may include an axially-oriented inlet 52 located at a first axial end 54 of turbocharger 36, and a tangentially-oriented volute 56 located between first axial end 54 and a second axial end 58 of turbocharger 36. Turbine shroud 44 may include a volute 60 located between volute 56 and second axial end 58 of turbocharger 36. Turbine shroud 44 may be configured to receive exhaust flow from exhaust manifold 34 in a tangential direction at a volute inlet (not shown). Volute 60 may direct the exhaust flow in three directions: axially (along rotation axis X), radially inward (along a radius of the volute), and tangentially (around a rotation axis X) toward and through a nozzle ring 62. Nozzle ring 62 may be disposed downstream of volute 60 and be configured to accelerate exhaust gas flowing therethrough.

As compressor wheel 46 is rotated, air may be drawn axially into turbocharger 36 via inlet 52 and directed toward compressor wheel 46. Blades 64 of compressor wheel 46 may then push the air radially outward in a spiraling fashion and into intake manifolds 30 (referring to FIG. 1) via an outlet volute (not shown). Similarly, as exhaust from exhaust system 16 is directed axially, radially, and tangentially inward toward turbine wheel 48, the exhaust may push against blades 66 of turbine wheel 48, causing turbine wheel 48 to rotate and drive compressor wheel 46 via shaft 50. After passing through turbine wheel 48, the exhaust flow may exit axially outward through a turbine outlet 68 located at second axial end 58 of turbocharger 36 into aftertreatment system 38 (shown only in FIG. 1).

As illustrated in FIG. 3, turbine wheel 48 may be generally disc-shaped and include a generally annular hub 70. Blades 66 may extend outward in three dimensions from annular hub 70. Nozzle ring 62 may be located radially upstream of turbine wheel 48 (i.e., at a periphery of turbine wheel 48). While turbine wheel 48 rotates in a rotational direction R, nozzle ring 62 may be stationary. Nozzle ring 62 may be generally ring-shaped, and include an inner annular hub 72 and an outer annular flange 74. A plurality of three-dimensional vanes 76 may be disposed between inner annular hub 72 and outer annular flange 74 to direct and accelerate exhaust flow from volute 60 toward blades 66 of turbine wheel 48.

As shown in FIG. 3, each blade 66 may include an airfoil 78 having a lower face (also known as a hub face) 80 that is connected to hub 70, an opposing upper face (also known as a shroud face) 82 that is oriented towards an inner surface of shroud 44, a trailing edge 84 that is proximate to turbine outlet 68, a leading edge 86 that is opposite to trailing edge 84, a high-pressure side (also known as the pressure side) 88, and an opposing low-pressure side (also known as the suction side) 90. It is contemplated that trailing edge 84 may be located closer to turbine outlet 68 than leading edge 86.

Similarly, each vane 76 may include a lower face (also known as a hub face) 92 that is connected to nozzle ring 62, an opposing upper face (also known as a shroud face) 94 that is oriented towards an inner surface of shroud 44, a trailing edge 96 located proximate to turbine wheel 48, a leading edge 98 that is opposite to trailing edge 96, a high-pressure side (also known as the pressure side) 100, and an opposing low-pressure side (also known as the suction side) 102. It is contemplated that trailing edge 96 may be located closer to turbine wheel 48 than leading edge 98,

FIG. 4 illustrates a side-view of turbine wheel 48. For the purposes of this disclosure, a blade forward sweep angle α_(B) of blade 66 may refer to an angle between leading edge 86 of blade 66 and a base of hub 70. A meridional length L_(MB) of blade 66 may refer to a meridional distance between trailing and leading edges 84, 86 of blades 66 along a camber line passing through a lengthwise center of the blades. Blades 66 may curve along their lengths, each forming a corresponding meridional blade angle β_(B) defined by the following equation:

${\tan \left( \beta_{B} \right)} = \frac{r\; {\theta}}{z_{m}}$

θ=Angular coordinate, polar angle, or wrap angle

z_(m)=Local meridional coordinate along the meridional length

r=Local radial location

β_(B)=Local meridional blade angle

A thickness T_(B) may refer to a distance between low- and high-pressure sides 88, 90 that is generally orthogonal to the camber line. A spacing S_(B) may refer to a straight line distance between adjacent trailing edges 84 of adjacent blades 66. A solidity ratio SR_(B) of blade 66 may be defined as the ratio of the meridional chord length L_(MB) to the spacing S_(B) (SR_(B)=L_(MB)/S_(B)).

FIG. 5 illustrates a meridional view of a single blade 66 taken along the meridional length L_(MB). In the meridional plane shown in FIG. 5, an R-axis defines a radial direction, and a Z-axis defines an axial direction along the meridional length. FIG. 5 shows an inlet flow passage 104 adjacent to leading edge 86 (i.e., where exhaust enters leading edge 86 of blade 66), and an outlet flow passage or diffuser 106 adjacent to trailing edge 84 (i.e., where exhaust exits trailing edge 84 of blade 66). FIG. 5 also shows a hub curve 108 corresponding to hub face 80 and a shroud curve 110 corresponding to shroud face 82, it should be noted that the relationship between hub face 80 and shroud face 82 possess unique geometric and “ruled element blade” characteristics. In a ruled element blade, an angular location is defined by a straight line drawn in 3D space between points at span locations along the hub and shroud faces 80, 82, it should also be noted that hub curve 108 and shroud curve 110 are master curves and control generation of all other defining curves (e.g., intermediate curves between hub curve 108 and shroud curve 110). Modification of the hub and/or shroud curves 108, 110 may result in a subsequent modification of the intermediate curves.

For the purposes of this disclosure, a blade inlet cone angle λ_(B) may refer to an angle between the R-axis of the meridional plane and leading edge 86 of blade 66. An inlet hub radius r_(4H) may refer to a distance from the Z-axis of the meridional plane to a point on the hub curve 108 at leading edge 86. An inlet shroud radius r_(4S) may refer to a distance from the Z-axis of the meridional plane to a point on the shroud curve 110 at leading edge 86. An inlet width W_(B) may refer to a distance between the point on the hub curve 108 at leading edge 86 and the point on the shroud curve 110 at leading edge 86. An inlet width ratio WR_(B) may be defined as the ratio of the width W_(B) to the meridional length L_(MB) (WR=W_(B)/L_(MB)) AZ-axis offset Z_(B) may refer to a distance between the R-axis and the point on the hub curve 108 at leading edge 86. A non-dimensional Z-axis offset ratio ZR may be defined as the ratio of the Z-axis offset Z_(B) to the meridional length L_(MB) (ZR_(B)=Z_(B)/L_(MB)). An exit deviation angle (or clip angle) δ_(B) may refer to an angle between trailing edge 84 of blade 66 and the R-axis of the meridional plane. An exit hub radius r_(5H) may refer to a distance from the Z-axis of the meridional plane to a point on the hub curve 108 at trailing edge 84. An exit shroud radius r_(SS) may refer to a distance from the Z-axis of the meridional plane to a point on the shroud curve 110 at trailing edge 84. A turbine trim TR_(B) may be defined by the following equation: [(r_(5X)/r_(4S))²×100)]. A diffuser hub exit radius r_(6R) TR_(B) may refer to a distance from the Z-axis of the meridional plane to a point on the hub curve 108 at the diffuser 106. A diffuser shroud exit radius r_(5s) may refer to a distance from the Z-axis of the meridional plane to a point on the shroud curve 110 at the diffuser 106.

The aerodynamic performance of a radial and mixed flow turbine is usually interpreted as a function of velocity ratio U/C₀, where U is the blade tip speed and C₀ is the isentropic velocity, resulting from ideal expansion of gas through a pressure ratio equal to that of the turbine. Since turbochargers often need to operate at low U/C₀ operating conditions (or high expansion ratio conditions at constant tip speed), there is a need for an efficient turbine stage design to operate at these low U/C₀ conditions with low aerodynamic losses (e.g., incidence loss). The disclosed geometry of blade 66 has been selected to provide a desired aerodynamic flow uniformity and guidance through turbine 32 that reduces flow misalignment (incidence) and results in improved performance and efficiency at wide operating conditions (especially at low U/C₀ conditions) of turbocharger 36. In addition, the disclosed geometry of blade 66 increases structural integrity and manufacturability of the blades. For example, each blade 66 may have a blade forward sweep angle α_(B) of about 25-55°. In one embodiment, the blade forward sweep angle α_(B) is about 47°. Blade 66 may also have a blade inlet cone angle λ_(B) of about 50-70°. In one embodiment, the blade inlet cone angle λ_(B) is about 58°. Blade 66 may further have a clip angle δ_(B) of about 0-14°. In one embodiment, the clip angle δ_(B) is about 7°. These angle ranges may help to reduce the incidence of exhaust flowing through turbine 32 and improve vibration characteristics of the turbine 32, thereby improving aerodynamic performance and structural integrity of turbocharger 36.

In the disclosed embodiment, the solidity ratio SR_(B) of blade 66 may be about 0.8-1.2, with about 10 to 17 blades 66 for a given turbine 32. In one embodiment, the solidity ratio SR_(B) is about 1.05 for a turbine 32 housing 13 blades. The turbine trim TR_(B) of blade 66 may be about 50-80. In one embodiment, the turbine trim TR_(B) is about 59. The width ratio WR_(B) of blade 66 may be about 0.2-0.42. In one embodiment, the width ratio WR_(B) is about 0.29. The Z-axis offset ratio ZR_(B), of blade 66 may be about 0.07-0.20. In one embodiment, the Z-axis offset ratio ZR_(B) is about 0.13. Each of these geometrical features may help to improve aerodynamic performance and structural integrity of blades 66, while at the same time allow for smooth curves that are conducive to improving manufacturability. In particular, these geometrical features may create a blade profile that is suitable for flank milling.

As described above, FIG. 5 shows an inlet flow passage 104 adjacent to leading edge 86 (i.e., where exhaust enters leading edge 86 of blade 66), and an outlet flow passage or diffuser 106 adjacent to trailing edge 84 (i.e., where exhaust exits trailing edge 84 of blade 66). The disclosed geometry of inlet flow passage 104 has been selected to provide a desired aerodynamic flow guidance and uniformity into the turbine blade leading edge 86 that also reduces flow misalignment (incidence) and results in improved performance and efficiency at wide operating conditions (especially at low INC, conditions) of turbocharger 36. The outlet flow passage 106 which acts like a diffuser 106 may have a diffuser ratio at hub (r_(5h)/r_(6h)) from 1.15 to 1.55. In one embodiment, diffuser ratio at hub is 1.35. The outlet flow passage 106 which acts like a diffuser 106 may have a diffuser ratio at shroud (r_(6s)/r_(5s)) from 1.02 to 1.10. In one embodiment, diffuser ratio at shroud is 1.07.

FIG. 6 shows an alternative embodiment of blade 66. In this embodiment, leading edge 86 of blade 66 is substantially concave rather than being substantially straight as shown in the embodiment of FIG. 5. It is contemplated that having a concave leading edge 86 may help to further improve flow alignment at wide operating conditions, in some applications.

In order to further improve manufacturability and aerodynamic performance of blades 66, the meridional blade angle β_(B) may change along the meridional length L_(MB). Specifically, FIG. 7 shows a plurality of curves corresponding to the meridional blade angle β₃ between hub curve 108 and shroud curve 110. It should be noted that each curve between hub curve 108 and shroud curve 110 may correspond with an intermediate layer of blade 66 between the hub and shroud faces 80, 82. As can be seen from a comparison of the plurality of curves, the meridional blade angle β₁₁ at hub face 80 may be generally larger than the meridional blade angle β_(B) at shroud face 82 (i.e., blade 66 may be more vertical at hub face 80). In addition, the meridional blade angle β_(B) at both faces reaches a maximum at leading edge 86 and a minimum at a trailing edge 84. In other words, the meridional blade angle β_(B) generally decreases from leading edge 86 to trailing edge 84. As also shown in FIG. 7, the meridional blade angle β_(B) may vary between about −5° and 30° at leading edge 86, and vary between about −40° and −80° at trailing edge 84. This blade angle distribution may help to reduce aerodynamic losses and, thus, improve performance and efficiency of turbocharger 36.

As shown in FIG. 8, the thickness T_(B) of blades 66 may also vary along their meridional length L_(MB). In particular, FIG. 8 shows a plurality of curves corresponding to the thickness T_(B) of blades 66 between the hub and shroud faces 80, 82 relative to the meridional length L_(MB) of blades 66. As can be seen from the plurality of curves, the thickness of blades 66 may reach a maximum thickness T_(Bmax) of about 10 mm at about 60-80% of the meridional length 140 and be thinnest at trailing and leading edges 84, 86. In one embodiment, the maximum thickness T_(Bmax) may be at about 68% of the meridional length L_(MB). Also shown in FIG. 8, the thickness of blades 66 may be substantially greater along hub face 80 than along shroud face 82. Finally, a maximum thickness at the leading edge 86 may be about 0.38×T_(Bmax), while a maximum thickness at the trailing edge 84 may be about 0.61×T_(Bmax). This smooth Bezier curve thickness distribution of blades 66 may improve the manufacturability of the blades, especially using flank milling processes, which can be lower in cost than alternative manufacturing processes.

Referring back to FIG. 3, exemplary disclosed geometry of vanes 76 of nozzle ring 62 will now be discussed. For the purposes of this disclosure, a meridional length L_(MV) may refer to a meridional distance between trailing and leading edges 96, 98 of vanes 76 along a camber line passing through a lengthwise center of the vanes. Similar to blades 66, vanes 76 may also curve along their lengths, each forming a corresponding meridional vane angle β_(V) defined by the following equation:

${\tan \left( \beta_{V} \right)} = \frac{r\; {\theta}}{z_{m}}$

θ=Angular coordinate, polar angle, or wrap angle

z_(m)=Local meridional coordinate along the meridional length

r=Local radial location

β_(V)=Local meridional vane angle

A thickness T_(V) may refer to a distance between high- and low-pressure sides 100, 102 that is generally orthogonal to the camber line of vane 76. A chord length L_(CV) may refer to a straight line distance between trailing and leading edges 96, 98 of vanes 76. A spacing S_(V) may refer to a straight line distance between adjacent trailing edges 96 of adjacent vanes 76. A solidity ratio SR_(V) may be defined as the ratio of the chord length L_(CV) to the spacing S_(V) (SR_(V)=L_(CV)/S_(V)). A width W_(V) may refer to a distance between hub face 92 and shroud face 94 at leading edge 86. A width ratio WR_(V) may be defined as the ratio of the width W_(V) to the chord length L_(CV) (WR_(V)=W_(V)/L_(CV)). A blade inlet shroud tip radius r₁ may refer to a distance from a center of turbine wheel 48 to leading edge 86 of blade 66 at shroud face 82. A vane leading edge radius r₂ may refer to a distance from the center of turbine wheel 48 to leading edge 98 of vane 76 at shroud face 82. A vane inlet radius ratio My may be defined as the ratio of the vane leading edge radius r₂ to the blade inlet shroud tip radius r₁. A nozzle inlet stagger angle φ_(V) may refer to an angle between the chord length L_(cv) and the vane leading edge radius r₂. A vane trailing edge radius r₃ may refer to a distance from the center of turbine wheel 48 to trailing edge 96 of vane 76 at shroud face 82. A vane exit radius ratio ER_(V) may be defined as the ratio of the vane trailing edge radius r₃ to the blade inlet shroud tip radius r₁.

Similar to blades 66, the disclosed geometry of vanes 76 has been selected to provide desired aerodynamic flow angles with improved flow uniformity at an exit of nozzle ring 62, increased structural integrity of the vanes, and low torque loading of the vanes 76. For example, each vane 76 may have a solidity ratio SR_(V) of about 0.7-1.2, with about 13 to 25 vanes 76 included around nozzle ring 62. In one embodiment, the solidity ratio is about 1.11, with 23 blades included around nozzle ring 62. The width ratio W_(RV) of vane 76 may be about 0.2-0.40. In one embodiment, the width ratio WR_(V) is about 0.23. The vane inlet radius ratio IR of vane 76 may be about 1.3-1.5. In one embodiment, the vane inlet radius ratio IR is about 1.36. The vane exit radius ratio ER of vane 76 may be about 1.05-1.3. In one embodiment, the vane inlet radius ratio ER is about 1.19. Finally, the nozzle inlet stagger angle φ_(V) of vane 76 may be about 60°-80°. In one embodiment, the nozzle inlet stagger angle φ_(V) is about 74°. Each of these geometrical features may help to reduce aerodynamic losses, reduce vane torque loading, and improve the structural integrity of vanes 76, while at the same time allow for smooth curves that are conducive to improving manufacturability.

Also, similar to blades 66, the meridional vane angle β_(V) of vanes 76 may change along the meridional length L_(MV). Specifically, FIGS. 9 and 10 show curves corresponding to the meridional vane angle β_(V) from leading edge 98 to trailing edge 96 for two different embodiments of nozzle ring 62. In each of these embodiments, the meridional vane angle β_(V) may vary in a range of about 50-80° along its meridional length.

In a first embodiment shown in FIG. 9, the meridional vane angle β_(V) may be substantially different along the hub and shroud faces 92, 94. Thus, two separate curves are shown. A hub curve 112 may correspond to the meridional vane angle β_(V) along the hub face 92, and a shroud curve 114 may correspond to the meridional vane angle β_(V) along the shroud face 94. In this embodiment, both curves 112, 114 may share a substantially S-shaped curve along the meridional length, representing a shape of the chamber of vane 76. However, the hub curve 112 may be substantially greater than the shroud curve 114 at each point along the meridional length between leading edge 98 and trailing edge 96.

In a second embodiment shown in FIG. 10, the meridional vane angle γ_(V) may be substantially equal along the hub and shroud faces 92, 94. Thus, only one curve is shown. FIG. 10 shows a curve 116 that corresponds to the meridional vane angle β_(V) for both the hub and shroud faces 92, 94. In this embodiment, vane 76 may also have a generally S-shaped camber. Further, in this embodiment, there may be an inclination angle from leading edge 98 to trailing edge 96, shown here as inclination curve 118. Both of the above embodiments of vanes 76 may be used with nozzle ring 62 depending on a desired application. Having two separate vane angle distributions for hub face 92 and shroud face 94 may help to improve vibratory′ response characteristics of turbine blades 66 by reducing High Cycle Fatigue strains at wide operating conditions. Having a single vane angle distribution for both the hub and shroud faces 92, 94 may be more suitable for improved stage aerodynamic performance at wide operating conditions.

As shown in FIG. 11, the thickness T_(V) of vanes 76 may vary along their meridional length L_(MV). In particular, FIG. 11 shows a curve 120 corresponding to the thickness T_(V) of vanes 76 relative to the meridional length L_(MV) of vane 76. As can be seen from the curve, the thickness of vanes 76 does not vary between hub and shroud faces 92, 94. The thickness may reach a maximum thickness of about 5.5 min at about 20-50% of the meridional length L_(MV) and be thinnest at trailing edge 96. In one embodiment, the thickness of blades 66 reaches a maximum at about 32% of the meridional length L_(my). Finally, maximum thickness at the leading edge may be about 0.25×T_(Vmax), while the maximum thickness at the trailing edge may be about 0.09×T_(Vn). In a similar manner to blades 66, this smooth thickness distribution of vanes 76 may improve the manufacturability of the vanes.

INDUSTRIAL APPLICABILITY

The disclosed turbocharger may be implemented into any power system application where charged air induction is utilized. In particular, the specific geometry, blade/airfoil angle, and thickness distribution of blades 66 and vanes 76 may result in overall lower aerodynamic losses and, thus, improved performance and efficiency of turbine 32. The uniform and well-guided flow exiting nozzle ring 62 may result in more uniform loading of nozzle ring 62 and turbine wheel 48. This may help to reduce cyclic loading on turbine wheel 48, extending the useful life of turbine wheel 48. Because exhaust flow may be substantially uniform and well-guided to each blade 66, mechanical and vibrational losses attributable to misaligned exhaust flow and turbine blade geometry may be significantly reduced. In addition, nozzle ring 62 and turbine wheel 48 may have low solidity as compared to an equivalent axial turbine stage and, thus, fewer vanes and blades. The reduction in vanes and blades may equate to a reduction in manufacturing costs. Finally, the smooth angle and thickness distribution of blades 66 and vanes 76 may allow these components to be manufactured using flank milling, which can be a cheaper alternative to other manufacturing processes.

It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed turbocharger. Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed turbocharger. It is intended that the specification and examples be considered as exemplary only, with a true scope being indicated by the following claims and their equivalents. 

What is claimed is:
 1. A turbine blade for a turbocharger, comprising: an airfoil having: a hub face connected to a turbine wheel hub of the turbocharger; a shroud face located opposite the hub face and oriented towards a turbine shroud of the turbocharger; a trailing edge; and a leading edge opposite the trailing edge, wherein an angle between the turbine wheel hub and the leading edge is about 25-55 degrees.
 2. The turbine blade of claim 1, wherein a solidity ratio of the turbine blade ranges from about 0.8 to 1.2 between the hub face and the shroud face.
 3. The turbine blade of claim 1, wherein an angle between a radial axis in a meridional plane of the turbine blade and the leading edge in the meridional plane is about 50-70 degrees.
 4. The turbine blade of claim 1, wherein an angle between a radial axis in a meridional plane of the turbine blade and the trailing edge in the meridional plane is about 0-14 degrees.
 5. The turbine blade of claim 1, wherein a ratio between an inlet width of the leading edge in a meridional plane of the turbine blade and a meriodional distance between the leading edge and the trailing edge ranges from about 0.20 to 0.42 between the hub face and the shroud face.
 6. The turbine blade of claim 1, wherein a ratio between a Z-axis offset in a meridional plane of the turbine blade and a meriodional distance between the leading edge and the trailing edge ranges from about 0.07 to 0.20 between the hub face and the shroud face.
 7. The turbine blade of claim 1, wherein the turbine blade has a blade angle that changes along its meridional length from about −80° to 30°.
 8. The turbine blade of claim 7, wherein the blade angle decreases from the hub face to the shroud face.
 9. The turbine blade of claim 7, wherein the blade angle is greatest for the turbine blade at the leading edge.
 10. The turbine blade of claim 9, wherein the blade angle ranges from about −5 to 30 degrees between the hub face and the shroud face at the leading edge.
 11. The turbine blade of claim 7, wherein the blade angle is lowest for the turbine blade at the trailing edge.
 12. The turbine blade of claim 11, wherein the blade angle ranges from about −40 to −80 degrees between the hub face and the shroud face at the trailing edge.
 13. The turbine blade of claim 1, wherein a thickness of the turbine blade varies along a meridional length of the turbine blade and decreases from the hub face to the shroud face.
 14. The turbine blade of claim 13, wherein the thickness is greatest at about 60-80% of the meridional length of the turbine blade.
 15. The turbine blade of claim 14, wherein a maximum thickness at the leading edge is about 0.38 times the greatest thickness along the meridional length.
 16. The turbine blade of claim 14, wherein a maximum thickness at the trailing edge is about 0.61 times the greatest thickness along the meridional length.
 17. The turbine blade of claim 1, wherein the leading edge is substantially straight.
 18. The turbine blade of claim 1, wherein the leading edge is substantially concave.
 19. A nozzle ring for a turbocharger, comprising: a ring-shaped generally flat plate having: an inner annular hub; and an outer annular flange radially spaced apart from the inner annular hub; and a plurality of vanes disposed between the inner annular hub and the outer annular flange, wherein a camber of each of the plurality of vanes is generally S-shaped along its meridional length from a leading edge to a trailing edge of each of the plurality of vanes.
 20. The nozzle ring of claim 19, wherein the S-shaped camber generally inclines along its meridional length from the leading edge to the trailing edge.
 21. The nozzle ring of claim 19, wherein a vane angle of each of the plurality of vanes changes along its meridional length in a range from about 50 to 80 degrees.
 22. The nozzle ring of claim 19, wherein: each of the plurality of vanes includes a hub face and a shroud face; and a vane angle along the hub face is substantially different from a vane angle along the shroud face.
 23. The nozzle ring of claim 19, wherein: each of the plurality of vanes includes a hub face and a shroud face; and a vane angle along the hub face is substantially equal to a vane angle along the shroud face.
 24. The nozzle ring of claim 19, wherein a thickness of each of the plurality of vanes varies along its meridional length, the thickness being greatest at about 20-50% of the meridional length.
 25. The nozzle ring of claim 24, wherein a maximum thickness at the leading edge is about 0.25 times the greatest thickness along the meridional length.
 26. The nozzle ring of claim 24, wherein a maximum thickness at the trailing edge is about 0.09 times the greatest thickness along the meridional length.
 27. The nozzle ring of claim 19, wherein: each of the plurality of vanes includes a hub face and a shroud face; and a solidity ratio of each of the plurality of vanes ranges from about 0.7 to 1.2 between the hub face and the shroud face.
 28. The nozzle ring of claim 27, wherein a ratio between a width of each of the plurality of vanes and a chord distance between the leading edge and the trailing edge ranges from about 0.20 to 0.40 between the hub face and the shroud face.
 29. The nozzle ring of claim 28, wherein a ratio between a distance from a center of a hub associated with the turbocharger to the leading edge of each of the plurality of vanes at the shroud face, and a distance from the center of the hub to a leading edge of a blade associated with the turbocharger at the shroud face ranges from about 1.3 to 1.5 between the hub face and the shroud face.
 30. The nozzle ring of claim 29, wherein a ratio between a distance from the center of the hub to the trailing edge of each of the plurality of vanes at the shroud face, and the distance from the center of the hub to the leading edge of the blade at the shroud face ranges from about 1.05 to 1.3 between the hub face and the shroud face.
 31. The nozzle ring of claim 30, wherein an angle between a line drawn between the center of the hub to the leading edge of each of the plurality of vanes at the shroud face, and a line drawn between the leading edge and the trailing edge is about 60-80 degrees.
 32. A turbocharger, comprising: a housing at least partially defining a compressor shroud and a turbine shroud; a compressor wheel disposed within the compressor shroud; a shaft connected to the compressor wheel; a turbine wheel disposed within the turbine shroud and connected to an end of the shall opposite the compressor wheel, the turbine wheel including: a generally annular hub; and a plurality of blades disposed radially around the annular hub, each of the plurality of blades including an airfoil having a hub face connected to the annular hub, a shroud face opposite the hub face and oriented towards the turbine shroud, a trailing edge, and a leading edge opposite the trailing edge, wherein an angle between a base of the annular hub and the leading edge is about 25-55 degrees; and a nozzle ring including: a ring-shaped generally flat plate located at a periphery of the turbine wheel; and a plurality of vanes disposed radially around an upper surface of the plate, wherein a camber of each of the plurality of vanes is generally S-shaped along its meridional length from a leading edge to a trailing edge of each of the plurality of vanes. 